This invention relates generally to gas turbine engines and, more particularly, to the method of controlling the radial clearance between rotating and non-rotating parts thereof during variable speed and temperature operating conditions.
The invention herein described was made in the course of or under a contract, or subcontract thereunder, with the United States Department of the Air Force.
In an effort to maintain a high degree of efficiency, manufacturers of turbine engines have strived to maintain the closest possible clearance between the engine rotor and the surrounding stator structure, since any gas which may pass therebetween represents a loss of energy in the system. If the system were to operate only under steady-state conditions, it would be a simple matter to establish the desired close clearance relationship between the rotor and stator to obtain the greatest possible efficiency without allowing frictional interference between the elements. However, in reality, all turbine engines must initially be brought from a standstill condition up to a steady-state speed, and then eventually decelerate to the standstill condition. This transitional operation is not compatible with the ideal low clearance condition just described since the variation in rotor speed also causes growth thereof because of mechanical expansion caused by centrifugal forces. The stationary stator, of course, does not grow mechanically and there is, therefore, relative mechanical growth between the two structures during periods of transitional operation. Further, as the turbine engine is brought up to speed from a standstill condition, the temperature of the gas passing therethrough is increased proportionately, thereby exposing both the rotor and the stator to variable temperature conditions. These conditions cause thermal growth of the two structures and, if the two structures have different thermal coefficients of expansion, which is generally true, then there is also the occurrence of relative thermal expansion between the elements. Characteristically, a rotor is necessarily a large mass element which allows it to rotate at very high speeds, thereby inherently yielding a very slow thermal response (high thermal inertia). On the other hand, the stator is a stationary element and, preferably, has a high thermal response (low thermal inertia) to allow for thermal growth of the stator during periods of acceleration to accommodate the mechanical growth of the rotor during those periods.
Early turbomachines were designed to operate at relatively low speeds and low temperatures. The stationary shrouds were supported by structures bathed in cold air, and thereby exhibited a minimal amount of thermal growth and a slow temperature response to transients. The relative clearance between the rotor and the shroud was therefore determined by the radial growth of the rotor structure. But, since compressor-discharge air temperatures in the engine were relatively cool, and the turbomachinery ran at relatively low speeds, rotor growth due to temperature and centrifugal loading was moderate and therefore not a problem. Thus, proper clearance control between the relatively rotating parts was not considered to be a significant problem.
As the technology developed, and a single stage turbine was introduced, the operational speeds of the rotor, and the discharge temperature of the compressor were significantly increased. The resulting increased radial growth of the rotor, caused by centrifugal loading and thermal expansion, necessitated an accommodating growth of the shroud in order to maintain the proper radial clearances between the two parts. In order to accomplish this it was necessary to remove the cold bath from the support structure of the stationary shroud and to instead expose it to higher temperatures which allowed it to grow along with the rotor.
Since the efficiency and the wear life of the rotor and shroud portions of a gas turbine engine can be best optimized by operation at a particular radial clearance, the normal practice is to design the machine such that the desired clearance exists during maximum speed, steady-state operating conditions. As a consequence, however, during other periods of operation such as during transient operation, the clearance is less than the predetermined desired clearance. In order to accommodate this phenomenon by providing adequate clearance control during transients, a shroud support structure was preferably composed of a low-alpha material (having a low thermal coefficient of expansion), which in turn provided the required large cold clearances. However, with the use of the low-alpha material, relatively high clearances existed during part-power performance to thereby bring about reduced performance.
With the advent of gas turbine engines having still higher speeds and operating temperatures, the preferred low-alpha materials were found to be inadequate since they were not strong enough at high operating temperatures to ensure safe operation. The need for higher strength at higher temperatures called for the use of nickel-base alloys, whose coefficient of thermal expansion was characteristically higher than that of previously used metals. The nickel-base alloys gave adequate clearance control during maximum operating conditions and at part-power conditions, but the cold clearances between the rotating and non-rotating structures were thus reduced. And, during certain periods of transient operation, the clearances were reduced such that there was frictional contact established between the moving and non-moving parts, thereby resulting in wear and reduction of engine performance and efficiency. As is well known in the art, clearance between the two elements is at a minimum during periods of operation when the engine is decelerated to part power and then rapidly accelerated thereafter (hot rotor burst), and it is therefore this clearance which establishes the critical criteria for the design of an aircraft jet engine.
The problems associated with the maintenance of proper clearance between the turbine rotor and shroud apply equally as well to other relatively rotating parts of turbomachine. For example, throughout the length of a turbine engine there are various seal arrangements interposed between the moving and stationary parts of the engine to reduce or substantially prevent the axial flow of a motive fluid in the annular chamber defined by the two members. These seals are commonly provided between the rotating and non-rotating parts of the turbine and are referred to as static turbine seals. Another common seal is that used between the aft part of the rotating compressor and the circumscribing stationary casing, and is commonly known as the CDP (compressor discharge pressure) seal. These seals, with their associated stationary and rotating parts, are susceptible to the same phenomena as that of the shroud discussed hereinabove and efficiency and wear can be a problem when operating over a variable range of speeds and temperatures.
It is, therefore, an object of the present invention to provide a turbine engine which operates efficiently with desirable clearances over a wide range of speeds and temperatures.
It is also an object of the present invention to provide a turbine shroud structure which affords close clearance control at maximum and part-power, steady-state operating conditions.
It is a further object of the present invention to provide a turbine shroud structure which affords large cool clearances during low temperature operation and acceptable clearances during transient operation.
Yet another object of the present invention is to provide in a gas turbine engine a method for controlling the radial clearance between the rotating and non-rotating parts thereof during variable speed and temperature operating conditions.
These objects and other features and advantages become more readily apparent upon reference to the following description when taken in conjunction with the appended drawings.